Typically, gas turbine engines include a compressor for compressing air, a combustor for mixing the compressed air with fuel and igniting the mixture, and a turbine blade assembly for producing power. Combustors often operate at high temperatures that may exceed 2,500 degrees Fahrenheit. Typical turbine combustor configurations expose turbine blade assemblies to these high temperatures. As a result, turbine blades and turbine vanes must be made of materials capable of withstanding such high temperatures. Turbine blades, vanes, transitions and other components often contain cooling systems for prolonging the life of these items and reducing the likelihood of failure as a result of excessive temperatures.
This invention is directed to a cooling system for a transition duct for routing a gas flow from a combustor to the first stage of a turbine section in a combustion turbine engine. In one embodiment, the transition duct may have a multi-panel outer wall formed from an inner panel having an inner surface that defines at least a portion of a hot gas path plenum and an intermediate panel positioned radially outward from the inner panel such that one or more cooling chambers is formed between the inner and intermediate panels. In another embodiment, the transition duct may include an inner panel, an intermediate panel and an outer panel. The inner, intermediary and outer panels may include one or more metering holes for passing cooling fluids between cooling chambers for cooling the panels. The intermediary and outer panels may be secured with an attachment system coupling the panels to the inner panel such that the intermediary and outer panels may move in-plane.
The cooling system may be configured to be usable with any turbine component in contact with the hot gas path of a turbine engine, such as a component defining the hot gas path of a turbine engine. One such component is a transition duct. The transition duct may be configured to route gas flow in a combustion turbine subsystem that includes a first stage blade array having a plurality of blades extending in a radial direction from a rotor assembly for rotation in a circumferential direction, said circumferential direction having a tangential direction component, an axis of the rotor assembly defining a longitudinal direction, and at least one combustor located longitudinally upstream of the first stage blade array and may be located radially outboard of the first stage blade array. The transition duct may include a transition duct body having an internal passage extending between an inlet and an outlet. The transition duct may be formed from a duct body that is formed at least in part from a multi-panel outer wall. The multi-panel outer wall may be formed from an inner panel having an inner surface that defines at least a portion of a hot gas path plenum and an intermediate panel positioned radially outward from the inner panel such that at least one cooling chamber is formed between the inner and intermediate panels. The multi-panel outer wall may also include an outer panel positioned radially outward from the intermediate panel such that at least one cooling chamber is formed between the intermediate and outer panels.
The cooling system may include one or more metering holes to control the flow of cooling fluids into the cooling chambers. In particular, the outer panel may include a plurality of metering holes. The intermediate panel may include one or more impingement holes, and the inner panel may include one or more film cooling holes.
The invention is also directed to a method of forming a multi-panel outer wall including an impingement cooling panel for components that are used under high thermally stressed conditions and having complex outer surface contours. The method comprises providing a component to be incorporated in a machine and perform in an environment of high thermally stressed conditions and having an inner panel having an outer surface with an array of interconnected ribs disposed on the outer surface. An intermediate panel is positioned over the component to cover at least a portion of the outer surface and ribs of the component.
The method also includes applying an external force under pressure across a surface area of the intermediate panel against the outer surface of the component to contour the intermediate panel according to a geometric configuration formed by the ribs. In performing this step the cooling chambers are formed between the outer surface and ribs of the component and the intermediate panel. In addition, the method may also comprise forming one or more holes in the intermediate panel and inner panel to allow airflow into and out of the cooling chambers.
The intermediate panel may then be affixed to the inner panel by known techniques. More specifically, the intermediate panels are affixed to the inner panel at first sections of the intermediate panel that contact the ribs on the inner panel.
The cooling system formed from a three-layered system is particularly beneficial for a transvane concept, where the hot gas flow is accelerated to a high Mach number, and the pressure drop across the wall is much higher than in traditional transition ducts. This high pressure drop is not ideal for film cooling, and an impingement panel alone is insufficient to reduce the post-impingement air pressure for ideal film cooling effectiveness. Therefore, the outer panel, which serves primarily as a pressure drop/flow metering device, is especially needed for this type of component.
Upstream portions of the transvane, where the hot gas path velocity is lower and the pressure difference across the wall is also lower, may benefit from the two wall construction, which is the embodiment with the outer wall including the metering holes or wherein the intermediate panel with the impingement holes are sufficient to drop the pressure for film effectiveness.
These and other embodiments are described in more detail below.